It is known to cast, forge or otherwise manufacture integrally bladed disks for use in compressor or turbine rotors in gas turbine engines especially for aircraft. Several techniques are known in the art for forming integrally bladed rotor disks, the present invention is adapted for use with those techniques in which the blades are formed in situ from solid as opposed to being formed separately and joined to the rotor. In the present context integrally bladed rotor disks shall be construed as including not only single stage disks (known as blisks) but also multi-stage assemblies either joined together or formed integrally, single stage integrally bladed rings (known as blings) and multi-stage bladed ring or drum assemblies (known as blums).
It is also well known to provide blades with abrasive tips so that if, or where, they come close to a surrounding portion of an engine casing the blade tips can cut a wear track in an encircling liner specially provided for the purpose. By this technique radial clearance gaps can be minimised to reduce overtip gas leakage and thereby improve engine efficiency. Suitable coating deposition methods for applying selectively a coating of a hard first material to a surface of a body comprised of a second, relatively softer material are known under the generic term of friction surfacing, and are described in British patent documents GB 2268430, GB 2242848, GB 2224682 and GB 2210572. However, other techniques of applying surface coating may be employed such as for example plasma spray coating.
Hitherto the hard or abrasive tips have been applied to blades individually as taught for example in British Patent Application 2153447. This is a time consuming and therefore expensive operation since each blade is treated separately. The present invention is intended to overcome this drawback by applying the hard tip facing to all blades in the assembly, in effect, simultaneously.